US4055044A - Rocket engine construction and connection for closed and opened fluid cooling circuits for the walls thereof - Google Patents

Rocket engine construction and connection for closed and opened fluid cooling circuits for the walls thereof Download PDF

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Publication number
US4055044A
US4055044A US05/522,990 US52299074A US4055044A US 4055044 A US4055044 A US 4055044A US 52299074 A US52299074 A US 52299074A US 4055044 A US4055044 A US 4055044A
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United States
Prior art keywords
section
cooling channels
inlet
cooling
discharge
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Expired - Lifetime
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US05/522,990
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English (en)
Inventor
Helmut Dederra
Gunther Schmidt
Jurgen Stanke
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Airbus Defence and Space GmbH
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Messerschmitt Bolkow Blohm AG
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F02K9/62—Combustion or thrust chambers
    • F02K9/64—Combustion or thrust chambers having cooling arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/97—Rocket nozzles
    • F02K9/972—Fluid cooling arrangements for nozzles

Definitions

  • This invention relates in general to the construction of rocket engines and, in particular, to a new and useful fluid cooled rocket engine combustion chamber having walls with cooling channels therein, which are divided longitudinally, and which includes at least one section wherein a propellant cooling fluid is circulated through cooling channels which are opened at their outer ends to discharge in the direction of thrust.
  • a main coolant circuit including passages through which the entire wall of the nozzle inclusive of the combustion chamber is cooled in countercurrent flow, and which includes a partly opened coolant circuit for the throat section which is advantageously designed as a replaceable insert.
  • the propellant volume of the main coolant circuit is supplied to the injector head of the combustion chamber.
  • the partial coolant volume of the open coolant circuit cools the wall of the throat section insert first regeneratively and, thereupon, by film cooling. Because of the pressure increase in the last portion of the coolant circuit, the regenerative cooling of the entire nozzle, including the combustion chamber, requires relatively thick walls of the cooling tubes.
  • a further disadvantage of the known constructions is particularly apparent when the rocket engines are operated in a vacuum condition.
  • the downstream final portion of the nozzle must be flared by an amount greater than the optimal flaring for ground-tested rocket nozzles. This results in an expensive construction, and correspondingly higher costs.
  • the engines equipped with nozzles for vacuum operation have to be operated in a vacuum chamber.
  • the present invention provides a rocket engine combustion chamber construction wherein the thrust nozzle comprises at least two independent cooling circuits in which the cooling is ensured by the formation of cooling flow structures of relatively light construction so that the coolant losses are low, and which are constructed so that they can be tested on the ground in an inexpensive manner.
  • an opened coolant circuit having its inlet closely adjacent the downstream inlet of the first or closed coolant circuit or circuits. With the construction, the opened coolant circuit cools the thrust nozzle wall which extends in the divergent nozzle section to the open nozzle discharge.
  • the invention provides the advantages of the known constructions without their disadvantages. If only one closed coolant circuit is provided, its inlet is located approximately in the middle of the divergent portion of the nozzle. Therefore, during its passage to the outlet which is located at the injection head, the coolant is heated up only to an extent which permits it to cope with the pressures which are produced while using usual materials and wall thicknesses for the cooling tubes.
  • the open coolant circuit cooling the downstream portion of the nozzle requires only thin tubes into which approximately 1% of the entire coolant quantity is supplied under a relatively low pressure to pass therethrough in parallel flow.
  • the high temperatures increase thus obtainable causes the propellant used as a coolant to leave the thin tubes at a high velocity producing an additional thrust, whereby, the propellant losses are reduced to extremely small quantities.
  • the heated up coolant of one of the circuits serves to drive auxiliary units.
  • the sectioning into at least three, where the open circuit is considered also, coolant circuits of which the second closed circuit may cool a thermally high stressed zone of the nozzle, offers the advantage of a further thermal relief of the main coolant quantity to be supplied to the injector head, and also the possibility of utilizing the propellant quantities to be used in the auxiliary units for regenerative cooling.
  • the open circuit tubes are provided with nozzle-shaped outlets and, in the preferred form, the nozzle outlets may comprise an annular portion forming an auxiliary thrust nozzle.
  • the open coolant circuit associated with the downstream final portion of the nozzle may, in accordance with another development of the invention, open into the interior of the thrust nozzle and the coolant thus introduced into the interior of the nozzle is used, in accordance with the invention, as a film coolant passing over the interior walls of the combustion chamber.
  • either the thrust nozzle in the entire zone of the open coolant circuit is designed as a self-contained structural part and detachably connected to the fore part of the thrust nozzle, or only the portion of the thrust nozzle extending rearwardly of the coolant inlet of the open coolant circuit is designed as a self-contained extension of the thrust nozzle.
  • the front portion of the nozzle extending upstream of the plane of the separation between the two parts is flared so that, even under atmospheric conditions on the ground, an optimal operation of the engine is ensured.
  • Only the detachable part of the nozzle is desigend with a divergence which is necessary for the vacuum operation.
  • Such a construction makes it possible, after dismounting the rear portion of the nozzle or the nozzle extension, to test vacuum nozzles in ground test stands and to measure their cooling properties under real conditions.
  • a rocket engine combustion chamber which includes a first combustion chamber wall portion having circumferentially arranged longitudinally extending cooling channels therethrough which are connected in a closed circuit through an annular inlet at one end which is connected to the propellant pump and an annular discharge at the opposite or closed combustion chamber end which is connected to the injector and which includes a second section having circumferentially arranged and longitudinally extending second cooling channels which are connected to an inlet duct at its inner end, and at its outer end, it is opened to discharge in the direction of thrust, either directly to the atmosphere or into the interior of the thrust engine at the downstream end of the divergent portion of the nozzle.
  • a further object of the invention is to provide an improved rocket engine combustion chamber construction, wherein the walls of the divergent section are made-up of a plurality of individual parts which are flanged together with the flange part of one having an annular inlet duct for the coolant which is connected through a passage to a first set of circumferentially arranged and longitudinally extending cooling channels which extend inwardly to the closed end of the combustion chamber and which includes a construction wherein a wall of the duct which is common to the flange connection between the two parts is provided with a relatively small nozzle flow passage from the inlet to a second set of coolant cooling channels in the trailing or downstream end of the divergent section of the nozzle, which terminate in open ends for discharge of a small portion of the coolant through the open ends in the direction of thrust.
  • a further object of the invention is to provide a rocket engine combustion chamber construction and to an improved flange construction between two sections of the rocket combustion chamber which are simple in design, rugged in construction and economical to manufacture.
  • FIG. 1 is a partial longitudinal sectional view of a rocket engine constructed in accordance with the invention
  • FIGS. 1a and 1b are detailed sectional views of alternate embodiments of the discharge end of the second cooling channels
  • FIG. 2 is a view similar to FIG. 1 of another embodiment of the invention.
  • FIG. 3 is a view similar to FIG. 1 of still another embodiment of the invention.
  • FIG. 4 is an enlarged partial sectional view of the details of construction of the flange connection between two cooling channel sections of a rocket engine combustion chamber.
  • the invention embodied therein comprises a rocket engine, generally designated 50, which includes a closed end or head 52 and a tubular wall part, generally designated 54, which is divided into sections 1 and 2.
  • Sections 1 and 2 are connected together in end-to-end relationship in order to form a combustion chamber, which includes a uniform diameter portion 1a, a converging portion 1b, a narrow neck portion or throat 1c and a divergent portion 1d.
  • Section 2 is entirely in a divergent portion of the combustion chamber which has an opened end or thrust gas discharge 56.
  • section 1 includes a plurality of circumferentially arranged longitudinally extending cooling channels 12 which are connected at their outer ends to an annular feed manifold 11.
  • the feed manifold 11 has an intake 11a which is connected, for example, to a supply tank for a propellant component through a pump (not shown).
  • the propellant component forms a coolant which is directed in the direction of arrow 58 through the inlet 11a and flows through each individual cooling channel 12 to an annular outlet manifold or duct 13.
  • Duct 13 includes a connection 13a for the outflow of the propellant in the direction of arrow 60 through a conduit 62 to an injection nozzle 64 where the preheated propellant component is directed into combustion chamber 1a in the form of a spray which is reacted within the combustion chamber.
  • At least a portion of the divergent section which includes both the entire section 2 and a portion 1d of section 1, is formed with cooling channels 22 which are circumferentially arranged and which extend generally longitudinally.
  • An annular inlet 21 is connected to direct a somewhat smaller portion of a propellant component through the individual cooling channels 22 which, at the thrust nozzle discharge side 56, are opened so as to form a discharge opening 23 for directing the coolant which will probably be in a vaporized condition in the direction of the arrow 62 which is in the same general direction as the thrust gas discharge of the rocket engine.
  • the inlet manifold 21 is provided with an inlet connection 21a for connecting it to a source of a propellant component which flows in the direction of the arrow 65.
  • the cooling channels 22 therefor are formed by part of an open coolant circuit which has as open discharge end 23.
  • the coolant passes at high velocity through cooling channels 22 and escapes again at the ends thereof thereby producing an additional impetus or thrust.
  • the end portions of tubes 22 are advantageously designed as individual nozzles 24, as shown in FIG. 1a or the ends of the individual cooling channels 22 are connected to an annular nozzle manifold 25 which provides an auxiliary thrust nozzle for supplying a uniform auxiliary thrust around the periphery of the main thrust discharge 56.
  • cooling channels 12 and 22 may be designed as straight longitudinal cooling channels or they may be wound in a spiral from their inlet ends to their discharge ends
  • section 2 is divided into section 2' and 3.
  • Section 2' includes the cooling channels 22' which form an open circuit and which have a discharge end 23' which permits a discharge of the coolant in the direction of arrow 62'.
  • the structural part 3 comprises an intermediate section which is cooled by a closed coolant circuit formed of longitudinally extending and circumferentially arranged coolant channels 32 which have an inlet manifold 31 adjacent the section 2' and a discharge manifold 32 adjacent the section 1'.
  • the coolant which flows through the intermediate coolant channels 32 may be employed for the actuation of auxiliary units, which are not shown.
  • the structural part 2' has a wider divergence than the section 2 of FIG. 1. Such a construction would be empolyed, for example, for operating the rocker engine in a vacuum.
  • the structural parts 2' and 3 are provided with flanges 34 and 36, respectively, which abut at a separation plane 35.
  • a rocket engine 50" includes sections 1", 2" and 4.
  • the section 2" corresponds substantially to section 2' shown in FIG. 2, but instead, it has an outer end with a flange 41 which abuts against, and is connected to, a flange 43 of an outer section 4.
  • the two flanges abut at a separation plane 43'.
  • Outer section 4 includes a single divergent outer wall 42, and this wall is cooled by directing the gases or vapors from coolant channel 22" in the direction of arrow 70, as shown in FIG. 3.
  • This outer wall 42 is more divergent than the structural part 2' of FIG. 2, and this is necessary for vacuum operation.
  • FIG. 4 shows the details of construction of the interconnection of sections 2' and 3 of a type similar to that shown in FIG. 2.
  • the connection may be employed for supplying coolant to both the closed and open circuits.
  • the supply or inlet assembly includes a housing, generally designated 70, forming an inlet flow passage which communciates with an annular manifold 72 which serves as a distributor for the coolant.
  • Manifold 72 is provided with bores 73 which individually communicate with cooling channels 32 of the closed coolant circuit of structural part 3.
  • Manifold 72 is also provided with further bores for control passages 74 in which calibrating nozzles 55 are provided and which serve to control the small quantity of coolant which will be permitted to flow from the manifold into an annular inlet duct 80 connected to the open coolant circuit through tubes 22'.
  • the connecting flanges 34 and 36 are provided with extensions 76 and 77 which are secured to each other by means of bolts 78.
  • a similar arrangement may also be provided for the embodiment shown in FIGS. 1 and 3.
  • FIGS. 1 through 4 are intended only for purposes of illustration, and the various sections shown in each embodiment may be used in another sequence or in distinct lengths which are different from that indicated and in accordance with the specifal requirements of the cooling system.
  • a particular coolant circuit may be provided for the thermally high stressed throat section 1c.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Testing Of Engines (AREA)
US05/522,990 1973-11-13 1974-11-11 Rocket engine construction and connection for closed and opened fluid cooling circuits for the walls thereof Expired - Lifetime US4055044A (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DT2356572 1973-11-13
DE2356572A DE2356572C3 (de) 1973-11-13 1973-11-13 Flüssigkeitsgekühlte Raketenbrennkammer mit Schubdüse

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US4055044A true US4055044A (en) 1977-10-25

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US05/522,990 Expired - Lifetime US4055044A (en) 1973-11-13 1974-11-11 Rocket engine construction and connection for closed and opened fluid cooling circuits for the walls thereof

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US (1) US4055044A (cs)
JP (1) JPS5823496B2 (cs)
DE (1) DE2356572C3 (cs)
FR (1) FR2250899B1 (cs)
IN (1) IN142651B (cs)

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4223530A (en) * 1977-09-30 1980-09-23 Erich Kirner Liquid fuel rocket engine having a propellant component pump turbine with a secondary thrust discharge and to a method of operating a liquid fuel rocket engine
US4369920A (en) * 1979-12-08 1983-01-25 Messerschmitt-Bolkow-Blohm Gmbh Arrangement to cool the thrust nozzle for a rocket engine
US5269132A (en) * 1992-10-29 1993-12-14 E-Systems, Inc. Method and apparatus for controlling infrared emissions
US5765360A (en) * 1995-02-17 1998-06-16 Daimler-Benz Aerospace Ag Process for cooling engine walls and wall structure for carrying out the process
US6220036B1 (en) * 1997-04-15 2001-04-24 Mitsubishi Heavy Industries, Ltd. Cooling structure for combustor tail pipes
WO2003052255A1 (en) * 2001-12-18 2003-06-26 Volvo Aero Corporation A component for being subjected to high thermal load during operation and a method for manufacturing such a component
WO2004051068A3 (en) * 2002-12-02 2004-08-12 Aerojet General Co Nozzle with spiral internal cooling channels
US20060032212A1 (en) * 2004-08-10 2006-02-16 The Boeing Company Lightweight rocket engine combustion chamber and associated method
US20060144959A1 (en) * 2003-03-04 2006-07-06 Hewitt Ross A Rocket engine chamber with layered internal wall channels
US20080034851A1 (en) * 2006-08-11 2008-02-14 Wyle Laboratories, Inc. Emission controlled engine exhaust static test stand
US20080134667A1 (en) * 2006-07-24 2008-06-12 Thomas Clayton Pavia Systems, methods and apparatus for propulsion
FR2923269A1 (fr) * 2007-11-06 2009-05-08 Cnes Epic Tuyere propulsive de detente d'un jet supersonique de gaz
US20090288390A1 (en) * 2008-05-23 2009-11-26 Thomas Clayton Pavia Simplified thrust chamber recirculating cooling system
KR100959793B1 (ko) 2007-12-31 2010-05-28 한국항공우주연구원 액체로켓엔진용 양방향 재생냉각 연소실 및 이의 모사시험장치
US20100229565A1 (en) * 2006-12-19 2010-09-16 Arne Boman Wall of a rocket engine
US20120060464A1 (en) * 2007-07-24 2012-03-15 James Robert Grote Systems, methods and apparatus for propulsion
CN104948347A (zh) * 2014-03-31 2015-09-30 北京航天动力研究所 一种具有均流功能的推力室集合器
CN113266492A (zh) * 2021-04-16 2021-08-17 北京星际荣耀空间科技股份有限公司 发动机推力室、火箭发动机、液体火箭
CN113530718A (zh) * 2021-08-31 2021-10-22 西安航天动力研究所 一种火箭发动机推力室热试用身部模块
US20220090562A1 (en) * 2020-09-21 2022-03-24 Arianegroup Gmbh Combustion chamber section with integral baffle and method of making a combustion chamber section
CN114483381A (zh) * 2021-12-21 2022-05-13 西北工业大学 一种使用组合动密封的变结构发动机喷管

Families Citing this family (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2519538C2 (de) * 1975-05-02 1982-12-30 Messerschmitt-Bölkow-Blohm GmbH, 8000 München Verbindung zwischen dem Kühlmittel-Zulaufring und der Brennkammer-oder Schubdüsenwand eines Raketentriebwerks
DE3535779C1 (en) * 1985-10-07 1987-04-09 Messerschmitt Boelkow Blohm Arrangement for the cooling of rocket engine walls
DE3618038A1 (de) * 1986-05-28 1987-12-03 Messerschmitt Boelkow Blohm Stuetzstruktur fuer fluessigkeitsgekuehlte expansionsduesen
DE3710984A1 (de) * 1987-04-01 1988-10-13 Messerschmitt Boelkow Blohm Vakuumduesen-auslassring
FR2639404B1 (fr) * 1988-11-21 1994-04-15 Propulsion Ste Europeenne Divergent de moteur-fusee a tuyere annulaire complementaire
DE4315256A1 (de) * 1993-05-07 1994-11-10 Mtu Muenchen Gmbh Einrichtung zur Verteilung sowie Zu- und Abführung eines Kühlmittels an einer Wand eines Turbo-, insbesondere Turbo-Staustrahltriebwerks
WO1997014875A1 (en) * 1995-10-17 1997-04-24 Westinghouse Electric Corporation Gas turbine regenerative cooled combustor
DE10141108B4 (de) * 2001-08-22 2005-06-30 Eads Space Transportation Gmbh Raketentriebwerk mit geschlossenen Triebwerkskreislauf mit modularer Zuführung der Turbinenabgase

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2935841A (en) * 1956-06-18 1960-05-10 Bell Aircraft Corp Thrust chamber with integrated cooling and structural members
US3048966A (en) * 1958-12-15 1962-08-14 Snecma Rocket propulsion method
US3267664A (en) * 1963-03-19 1966-08-23 North American Aviation Inc Method of and device for cooling
US3303654A (en) * 1964-01-29 1967-02-14 Bringer Heinz Combustion chamber for ram-jets or rocket power units employing a cooling film of liquid fuel
GB1089055A (en) * 1965-07-20 1967-11-01 Bristol Siddeley Engines Ltd Combined combustion chamber and propulsive unit for a rocket engine
US3595023A (en) * 1967-01-16 1971-07-27 Bolkow Gmbh Rocket engine combustion chamber cooling
US3597923A (en) * 1969-10-02 1971-08-10 Michael Simon Rocket propulsion system
US3605412A (en) * 1968-07-09 1971-09-20 Bolkow Gmbh Fluid cooled thrust nozzle for a rocket

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2935841A (en) * 1956-06-18 1960-05-10 Bell Aircraft Corp Thrust chamber with integrated cooling and structural members
US3048966A (en) * 1958-12-15 1962-08-14 Snecma Rocket propulsion method
US3267664A (en) * 1963-03-19 1966-08-23 North American Aviation Inc Method of and device for cooling
US3303654A (en) * 1964-01-29 1967-02-14 Bringer Heinz Combustion chamber for ram-jets or rocket power units employing a cooling film of liquid fuel
GB1089055A (en) * 1965-07-20 1967-11-01 Bristol Siddeley Engines Ltd Combined combustion chamber and propulsive unit for a rocket engine
US3595023A (en) * 1967-01-16 1971-07-27 Bolkow Gmbh Rocket engine combustion chamber cooling
US3605412A (en) * 1968-07-09 1971-09-20 Bolkow Gmbh Fluid cooled thrust nozzle for a rocket
US3597923A (en) * 1969-10-02 1971-08-10 Michael Simon Rocket propulsion system

Cited By (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4223530A (en) * 1977-09-30 1980-09-23 Erich Kirner Liquid fuel rocket engine having a propellant component pump turbine with a secondary thrust discharge and to a method of operating a liquid fuel rocket engine
US4369920A (en) * 1979-12-08 1983-01-25 Messerschmitt-Bolkow-Blohm Gmbh Arrangement to cool the thrust nozzle for a rocket engine
US5269132A (en) * 1992-10-29 1993-12-14 E-Systems, Inc. Method and apparatus for controlling infrared emissions
US5765360A (en) * 1995-02-17 1998-06-16 Daimler-Benz Aerospace Ag Process for cooling engine walls and wall structure for carrying out the process
US6220036B1 (en) * 1997-04-15 2001-04-24 Mitsubishi Heavy Industries, Ltd. Cooling structure for combustor tail pipes
US20050188678A1 (en) * 2001-12-18 2005-09-01 Volvo Aero Corporation Component for being subjected to high thermal load during operation and a method for manufacturing such a component
WO2003052255A1 (en) * 2001-12-18 2003-06-26 Volvo Aero Corporation A component for being subjected to high thermal load during operation and a method for manufacturing such a component
US7299622B2 (en) 2001-12-18 2007-11-27 Volvo Aero Corporation Component for being subjected to high thermal load during operation and a method for manufacturing such a component
WO2004051068A3 (en) * 2002-12-02 2004-08-12 Aerojet General Co Nozzle with spiral internal cooling channels
US20040168428A1 (en) * 2002-12-02 2004-09-02 Aerojet-General Corporation Nozzle with spiral internal cooling channels
US6802179B2 (en) * 2002-12-02 2004-10-12 Aerojet-General Corporation Nozzle with spiral internal cooling channels
US7343732B2 (en) * 2003-03-04 2008-03-18 Aerojet-General Corporation Rocket engine chamber with layered internal wall channels
US20060144959A1 (en) * 2003-03-04 2006-07-06 Hewitt Ross A Rocket engine chamber with layered internal wall channels
US20060032212A1 (en) * 2004-08-10 2006-02-16 The Boeing Company Lightweight rocket engine combustion chamber and associated method
US20080134667A1 (en) * 2006-07-24 2008-06-12 Thomas Clayton Pavia Systems, methods and apparatus for propulsion
US7997510B2 (en) * 2006-07-24 2011-08-16 Thomas Clayton Pavia Systems, methods and apparatus for propulsion
US20080034851A1 (en) * 2006-08-11 2008-02-14 Wyle Laboratories, Inc. Emission controlled engine exhaust static test stand
US7793506B2 (en) 2006-08-11 2010-09-14 Wyle Laboratories, Inc. Emission controlled engine exhaust static test stand
US20100229565A1 (en) * 2006-12-19 2010-09-16 Arne Boman Wall of a rocket engine
US20120060464A1 (en) * 2007-07-24 2012-03-15 James Robert Grote Systems, methods and apparatus for propulsion
FR2923269A1 (fr) * 2007-11-06 2009-05-08 Cnes Epic Tuyere propulsive de detente d'un jet supersonique de gaz
WO2009068791A3 (fr) * 2007-11-06 2009-08-06 Centre Nat Etd Spatiales Tuyere propulsive de detente d ' un jet supersonique de gaz avec un flux secondaire d ' ecran
KR100959793B1 (ko) 2007-12-31 2010-05-28 한국항공우주연구원 액체로켓엔진용 양방향 재생냉각 연소실 및 이의 모사시험장치
US20090288390A1 (en) * 2008-05-23 2009-11-26 Thomas Clayton Pavia Simplified thrust chamber recirculating cooling system
CN104948347A (zh) * 2014-03-31 2015-09-30 北京航天动力研究所 一种具有均流功能的推力室集合器
US20220090562A1 (en) * 2020-09-21 2022-03-24 Arianegroup Gmbh Combustion chamber section with integral baffle and method of making a combustion chamber section
US11846255B2 (en) * 2020-09-21 2023-12-19 Arianegroup Gmbh Combustion chamber section with integral baffle and method of making a combustion chamber section
CN113266492A (zh) * 2021-04-16 2021-08-17 北京星际荣耀空间科技股份有限公司 发动机推力室、火箭发动机、液体火箭
CN113266492B (zh) * 2021-04-16 2022-03-15 北京星际荣耀空间科技股份有限公司 发动机推力室、火箭发动机、液体火箭
CN113530718A (zh) * 2021-08-31 2021-10-22 西安航天动力研究所 一种火箭发动机推力室热试用身部模块
CN114483381A (zh) * 2021-12-21 2022-05-13 西北工业大学 一种使用组合动密封的变结构发动机喷管
CN114483381B (zh) * 2021-12-21 2023-08-11 西北工业大学 一种使用组合动密封的变结构发动机喷管

Also Published As

Publication number Publication date
IN142651B (cs) 1977-08-06
FR2250899B1 (cs) 1980-12-26
DE2356572A1 (de) 1975-05-15
FR2250899A1 (cs) 1975-06-06
JPS5079614A (cs) 1975-06-28
JPS5823496B2 (ja) 1983-05-16
DE2356572C3 (de) 1979-03-29
DE2356572B2 (de) 1978-08-10

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